The mean camber line of a NACA 4-digit aerofoil is defined as: m y(x) = (2px x) p2 - x p m y(x) = = - (1 - p) : [(1 2p) + 2px x] - x > p where m defines maximum camber (percentage of the chord), p defines the position of the maximum camber (tenths of the chord), t defines the maximum thickness (percentage of the chord). Assume the chord c = 1 Consider a NACA 4412 aerofoil (m = 0.04, p = 0.4) as shown in the figure below. Use 2D Thin aerofoil theory to calculate the following: 1. The lift coefficient at zero angle of incidence 2. The no-lift incidence (the angle of incidence where no lift is generated) Figure Q1: NACA 4412 Aerofoil Note the following: T 1 k sin 0 - U (a A) U A cos(n) = (cos 0 = cos 01) - n=1 : 2T -de 2 T Ao = dy do An A. - = dx * dzy cos(no) do dx